Total temperature and total pressure measuring method for high Mach number aircraft

Abstract

The invention provides a total temperature and total pressure measurement method for a high-Mach-number aircraft, and the method comprises the steps: building a corresponding relation between wall surface pressure and total pressure through innovative molded surface design and structural design; obtaining a theoretical incoming flow total enthalpy and a theoretical incoming flow total temperature based on a sound velocity flow relationship; the theoretical incoming flow total enthalpy and the theoretical incoming flow total temperature are corrected by simulating a high-speed incoming flow environment in an arc wind tunnel. Based on the understanding of a high-Mach-number flow field mechanism, a high-Mach-number total temperature real-time measurement scheme is formed by solving the problem that the total pressure under the flight condition is difficult to integrally measure, and the method can be suitable for total temperature and total pressure measurement of a high-Mach-number aircraft.

Description

Technical Field

[0001] The invention belongs to the technical field of high-speed aircraft measurement, and in particular relates to a total temperature and total pressure measurement method for a high Mach number aircraft. Background Art

[0002] During the flight of a high-speed aircraft (with a Mach number exceeding 1), it is necessary to sense the total temperature and total pressure parameters of the incoming flow environment in real time to provide input for the aircraft's scramjet engine and precision guidance control. The measurement deviation of the total temperature and total pressure of the incoming flow directly affects the control of the aircraft.

[0003] At present, the measurement scheme of total pressure is relatively mature, while the measurement of total temperature is relatively complicated. Most of the total temperature measurements at home and abroad are aimed at ground wind tunnel conditions, and the total temperature is mainly achieved by measuring the total enthalpy, including the balanced sonic flow method, Fay-Riddell method, TDLAS total temperature measurement method and total enthalpy probe method. Among them, the balanced sonic flow method assumes that the airflow parameters before the nozzle throat are in a balanced, isentropic and steady state. According to the continuity equation, energy equation and thermodynamic properties of high-temperature gas, the relationship between the gas flow rate and total pressure, the cross-sectional area of ​​the sonic throat and the total enthalpy under different states is derived. This is a enthalpy measurement method currently used in aerodynamic thermal ground simulation tests, but the average volume enthalpy of the airflow before the nozzle throat is obtained by this method, and the total enthalpy of the airflow at a certain point in the flow field at the nozzle outlet cannot be obtained. In recent years, scholars have developed a water-cooled double-throat total enthalpy probe, which samples gas, measures the total pressure at the probe inlet (incoming flow total pressure) and the total temperature and total pressure before the second throat (used to calculate flow), and then combines the cross-sectional areas of the two throats to calculate the incoming flow total enthalpy and related parameters using the corresponding basic aerodynamic equations. However, due to the need for water cooling, it is not suitable for flight test measurements.

[0004] In general, the existing enthalpy measurement technology is mainly used in the flow field calibration of experimental equipment such as arc wind tunnels, and is concentrated in the field of non-contact measurement or requires water cooling to cool the flow to a measurable temperature, so it is not suitable for flight test measurements.

[0005] However, the existing technology lacks technical research on total temperature measurement at high Mach numbers in flight tests. When a high-speed aircraft flies at high Mach numbers, the total temperature of the incoming airflow is high. Currently, there is no effective method for measuring total temperature and total pressure at high Mach numbers suitable for flight tests. The main difficulties are as follows:

[0006] 1. When the aircraft flies at a high Mach number, the total temperature of the incoming flow is high. The commonly used non-contact temperature measurement method is not suitable for flight test measurement. The commonly used direct thermocouple can measure the highest temperature, which is far from meeting the measurement requirements of the total temperature of the incoming flow: for example, patent CN201921908182.7 provides an integrated temperature and pressure sensor assembly and CN201521064275.8 provides a composite total temperature measurement device, etc., both of which measure the total temperature and total pressure at room temperature and cannot be applied to the measurement of the working environment of high-speed aircraft;

[0007] 2. When the aircraft flies at a high Mach number, the leading edge of the head faces a harsh environment of high heat, high pressure, and strong scouring. The structural thermal protection for total temperature and total pressure measurement faces huge challenges, and technical research is needed;

[0008] 3. Generally, the total temperature measurement is placed at the leading edge of the aircraft, overlapping with the total pressure measurement. If they are measured separately, the head size will be too large, affecting the radar wave transmission, and thus failing to meet the terminal guidance requirements.

[0009] In summary, there is an urgent need to develop a total temperature and total pressure measurement method suitable for high Mach number aircraft to solve the problem of accurate measurement of total temperature and total pressure at high Mach number. Summary of the invention

[0010] The present invention aims to solve at least one of the technical problems existing in the prior art.

[0011] The present invention provides a total temperature and total pressure measurement method for a high Mach number aircraft, the method comprising:

[0012] According to the shape of the aircraft head, the double-throat nozzle profile design is carried out;

[0013] Design the cooling length of the double-throat nozzle according to the cooling range of the aircraft and the total airflow temperature;

[0014] A first opening point is set before the first throat, air is introduced through the opening, and the head wall pressure is measured; the head wall pressure is corrected to obtain the total pressure before the first throat;

[0015] A second opening point is set before the second throat, air is introduced through the opening, and the wall pressure is measured; the wall pressure is corrected to obtain the total pressure before the second throat;

[0016] When the airflow before the second throat is cooled to the measuring range of the temperature sensor, the temperature sensor is used to measure the total temperature of the airflow before the second throat;

[0017] Based on the sonic flow relationship, the theoretical total enthalpy of the incoming flow is obtained according to the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the geometric area of ​​the double nozzle throat;

[0018] Obtain theoretical total temperature of incoming flow according to theoretical total enthalpy of incoming flow;

[0019] In an arc wind tunnel, a high-speed incoming flow environment is simulated, and the measured incoming flow total temperature is obtained by using enthalpy measurement technology, and the measured incoming flow total enthalpy is calculated; the actual area ratio of the double nozzle throat is obtained according to the measured incoming flow total enthalpy, and the theoretical incoming flow total enthalpy calculation formula and the theoretical incoming flow total temperature calculation formula are corrected to obtain the corrected incoming flow total enthalpy calculation formula and the corrected incoming flow total temperature calculation formula;

[0020] Based on the revised incoming flow total enthalpy calculation formula and the revised incoming flow total temperature calculation formula, the total temperature and total pressure measurement of the high Mach number aircraft is realized.

[0021] The technical solution of the present invention provides a total temperature and total pressure measurement method for high Mach number aircraft, which establishes the corresponding relationship between wall pressure and total pressure through innovative surface design and structural design; obtains theoretical incoming flow total enthalpy and theoretical incoming flow total temperature based on the sound speed flow relationship; and corrects the theoretical incoming flow total enthalpy and theoretical incoming flow total temperature by simulating a high-speed incoming flow environment in an arc wind tunnel. The present invention is based on the understanding of the flow field mechanism of high Mach numbers, solves the problem that the total pressure under flight conditions is difficult to measure in an integrated manner, and forms a real-time total temperature measurement solution for high Mach numbers, which can be applied to the total temperature and total pressure measurement of high Mach number aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] The included drawings are used to provide a further understanding of the embodiments of the present invention, which constitute a part of the specification, are used to illustrate the embodiments of the present invention, and together with the text description, explain the principles of the present invention. Obviously, the drawings in the following description are only some embodiments of the present invention, and for ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0023] Figure 1 A schematic flow chart of a total temperature and total pressure measurement method for a high Mach number aircraft provided according to a specific embodiment of the present invention is shown;

[0024] Figure 2 A schematic cross-sectional view of a double-throat nozzle provided in a specific embodiment of the present invention is shown;

[0025] Figure 3 A schematic diagram of a double-throat nozzle profile according to a specific embodiment of the present invention is shown;

[0026] Figure 4 A schematic diagram of sensor positions provided according to a specific embodiment of the present invention is shown;

[0027] Figure 5 A schematic diagram of the Laval nozzle principle is shown. DETAILED DESCRIPTION

[0028] It should be noted that, in the absence of conflict, the embodiments in this application and the features in the embodiments can be combined with each other. The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. The following description of at least one exemplary embodiment is actually only illustrative and is by no means intended to limit the present invention and its application or use. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0029] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, it indicates the presence of features, steps, operations, devices, components and / or combinations thereof.

[0030] Unless otherwise specifically stated, the relative arrangement, numerical expressions and numerical values ​​of the parts and steps set forth in these embodiments do not limit the scope of the present invention. At the same time, it should be understood that, for ease of description, the sizes of the various parts shown in the accompanying drawings are not drawn according to actual proportional relationships. The technology, methods and equipment known to those of ordinary skill in the relevant art may not be discussed in detail, but in appropriate cases, the technology, methods and equipment should be considered as a part of the specification. In all examples shown and discussed here, any specific value should be interpreted as being merely exemplary, rather than as a limitation. Therefore, other examples of exemplary embodiments may have different values.

[0031] like Figure 1 As shown, according to a specific embodiment of the present invention, a total temperature and total pressure measurement method for a high Mach number aircraft is provided, the method comprising:

[0032] According to the shape of the aircraft head, the double-throat nozzle profile design is carried out;

[0033] Design the cooling length of the double-throat nozzle according to the cooling range of the aircraft and the total airflow temperature;

[0034] A first opening point is set before the first throat, air is introduced through the opening, and the head wall pressure is measured; the head wall pressure is corrected to obtain the total pressure before the first throat;

[0035] A second opening point is set before the second throat, air is introduced through the opening, and the wall pressure is measured; the wall pressure is corrected to obtain the total pressure before the second throat;

[0036] When the airflow before the second throat is cooled to the measuring range of the temperature sensor, the temperature sensor is used to measure the total temperature of the airflow before the second throat;

[0037] Based on the sonic flow relationship, the theoretical total enthalpy of the incoming flow is obtained according to the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the geometric area of ​​the double nozzle throat;

[0038] Obtain theoretical total temperature of incoming flow according to theoretical total enthalpy of incoming flow;

[0039] In an arc wind tunnel, a high-speed incoming flow environment is simulated, and the measured incoming flow total temperature is obtained by using enthalpy measurement technology, and the measured incoming flow total enthalpy is calculated; the actual area ratio of the double nozzle throat is obtained according to the measured incoming flow total enthalpy, and the theoretical incoming flow total enthalpy calculation formula and the theoretical incoming flow total temperature calculation formula are corrected to obtain the corrected incoming flow total enthalpy calculation formula and the corrected incoming flow total temperature calculation formula;

[0040] Based on the revised incoming flow total enthalpy calculation formula and the revised incoming flow total temperature calculation formula, the total temperature and total pressure measurement of the high Mach number aircraft is realized.

[0041] By applying this configuration, a total temperature and total pressure measurement method for high Mach number aircraft is provided. The corresponding relationship between wall pressure and total pressure is established through innovative surface design and structural design; the theoretical total enthalpy and theoretical total temperature of incoming flow are obtained based on the sound velocity flow relationship; and the high-speed incoming flow environment is simulated in an arc wind tunnel to correct the theoretical total enthalpy calculation formula and the theoretical total temperature calculation formula of incoming flow. The present invention is based on the understanding of the flow field mechanism of high Mach number, solves the problem that the total pressure under flight conditions is difficult to measure in an integrated manner, and forms a real-time total temperature measurement scheme for high Mach number, which can be applied to the total temperature and total pressure measurement of high Mach number aircraft.

[0042] For high-speed aircraft, the incoming flow has a high Mach number and a high total temperature, and conventional thermocouples cannot directly measure the temperature. Based on this, the present invention proposes a method of using airflow to stagnate and decelerate through double throats, while radiating heat to cool down, to obtain the total temperature of the airflow after cooling, and then obtain the total temperature of the airflow through aerodynamic simulation principles.

[0043] The Laval nozzle (supersonic nozzle) was invented by the Swede Laval in 1889. It is an important component for converting subsonic airflow into supersonic airflow. The schematic diagram is shown below. Figure 5 The Laval nozzle consists of two parts: a contraction section and an expansion section. During the flow inside the nozzle, the airflow converts thermal enthalpy energy into kinetic energy. The subsonic airflow passes through the contraction section and is accelerated to the speed of sound at the throat. It then continues to accelerate to the required Mach number in the expansion section, ultimately obtaining a supersonic airflow.

[0044] Based on this, in the present invention, a double-throat nozzle profile design is first carried out according to the shape of the aircraft head.

[0045] The double-throat nozzle is arranged at the head of the aircraft, such as Figure 2 As shown, the double Laval nozzle profile design is first carried out, and the flow field calculation is carried out. By adjusting the area ratio and the outlet position, the airflow can flow smoothly and reach the speed of sound in both throats. The first throat and the second throat are arranged in sequence in the incoming flow direction.

[0046] Preferably, the outlet position is usually selected at a position with lower back pressure, which can make the air flow smooth.

[0047] Preferably, to avoid the structure affecting radar wave transmission, ensure that the nozzle inlet and outlet diameters are no greater than 20 mm.

[0048] Furthermore, in the present invention, the cooling length of the double-throat nozzle is designed according to the temperature reduction range of the aircraft and the total airflow temperature.

[0049] Based on the total temperature of the incoming air from the aircraft and the maximum measurable range of the temperature sensor, the extent to which the total temperature of the gas needs to be reduced is determined, and aerodynamic heat / heat transfer coupling calculations are carried out so that the airflow is cooled by sufficient radiation heat dissipation and the total temperature drops to a temperature range that can be directly measured.

[0050] Furthermore, in the present invention, a first opening point is set before the first throat, the opening is used to bleed air, and the head wall pressure is measured; the head wall pressure is corrected to obtain the total pressure before the first throat.

[0051] Preferably, in order to make the pressure on the wall of the measuring head as close to the total pressure as possible, the position of the first opening point should be less than 10 mm from the front edge of the head.

[0052] By carrying out flow field simulation, the corresponding relationship between the head wall pressure at the first opening point and the total pressure before the first throat is established, and a first correction database is established; the total pressure before the first throat is obtained according to the first correction database.

[0053] Furthermore, in the present invention, a second opening point is set before the second throat, the opening is used to bleed air, and the wall pressure is measured; the wall pressure is corrected to obtain the total pressure before the second throat.

[0054] Preferably, in order to make the measured wall pressure as close to the total pressure as possible, the second opening point should be selected at a location with a lower airflow velocity according to the numerical simulation results.

[0055] By carrying out flow field simulation, the corresponding relationship between the wall pressure at the second opening point and the total pressure before the second throat is established, and a second correction database is established; the total pressure before the second throat is obtained according to the second correction database.

[0056] Further, in the present invention, when the airflow before the second throat is cooled to the measurement range of the temperature sensor, the temperature sensor is used to measure the total temperature of the airflow before the second throat.

[0057] Preferably, the temperature sensor may be a thermocouple.

[0058] In addition, if Figure 4 As shown, in order to ensure that the total temperature of the incoming flow before the second throat and the total pressure of the incoming flow before the second throat are measured at the same position, the center of the thermocouple node and the center of the second opening point should be located in the same plane and ensure insulation from the inner wall of the nozzle.

[0059] Furthermore, in the present invention, based on the sonic flow relationship, the theoretical total enthalpy of the incoming flow is obtained according to the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the throat area of ​​the double nozzle.

[0060] During the flow process, it is assumed that the gas performs a one-dimensional steady isentropic equilibrium flow. According to the basic flow dynamics equation, the theoretical total enthalpy of the incoming flow can be obtained through the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the throat area of ​​the double nozzle obtained in the previous steps.

[0061] Preferably, when the total enthalpy of the incoming flow is <2.326×10 6 J / kg, use formula (1) to calculate the theoretical total enthalpy of the flow,

[0062]

[0063] Where, h0 is the theoretical total enthalpy of the incoming flow, A1 and A2 are the geometric areas of the first throat and the second throat, respectively.

[0064] When the total enthalpy of the incoming flow is > 2.326×10 6 J / kg, use formula (2) to calculate the theoretical total enthalpy of the flow,

[0065]

[0066] Furthermore, in the present invention, the theoretical total temperature of the incoming flow is obtained by calculating the theoretical total enthalpy of the incoming flow.

[0067] Preferably, the theoretical total temperature of the incoming flow is calculated using formula (3):

[0068]

[0069] Among them, c p is the specific heat of air and T is the temperature.

[0070] Furthermore, in the present invention, in an arc wind tunnel, a high-speed incoming flow environment is simulated, and the enthalpy measurement technology is used to obtain the measured total temperature of the incoming flow, and the measured total enthalpy of the incoming flow is calculated; the actual area ratio of the double nozzle throat is obtained according to the measured total enthalpy of the incoming flow, thereby obtaining a corrected calculation formula for the total enthalpy of the incoming flow and a corrected calculation formula for the total temperature of the incoming flow.

[0071] Due to the influence of friction heat transfer loss and boundary layer, the actual throat area is lower than the throat geometric area, resulting in deviations in the theoretical total enthalpy calculation formula and the theoretical total temperature calculation formula. In the present invention, the measured total enthalpy of the incoming flow can be obtained according to the measured total temperature of the incoming flow, and the actual area ratio of the double nozzle throat can be obtained according to the measured total enthalpy of the incoming flow; based on the actual area ratio of the double nozzle throat, the revised total enthalpy calculation formula and the revised total temperature calculation formula of the incoming flow can be obtained.

[0072] Preferably, the enthalpy measurement technology used is the equilibrium sonic flow method, the Fay-Riddell method, the TDLAS total temperature measurement method or the total enthalpy probe method.

[0073] In addition, after the measured total temperature of the incoming flow is obtained by using the enthalpy measurement technology, the measured total enthalpy of the incoming flow is calculated according to formula (4):

[0074]

[0075] in, To measure the total temperature of the incoming flow, it is obtained based on the arc wind tunnel test.

[0076] Formula (5) can be used to obtain the actual area ratio of the double nozzle throat:

[0077]

[0078] According to the actual area ratio of the double nozzle throat The corrected total enthalpy h0′ calculation formula can be obtained:

[0079]

[0080] Furthermore, according to formula (7), the corrected calculation formula for the total incoming flow temperature T0′ can be obtained.

[0081]

[0082] Furthermore, after the correction is completed, formula (6) and formula (7) are directly used to carry out the total temperature and total pressure measurement of the high-speed aircraft.

[0083] The total temperature and total pressure measurement method for a high Mach number aircraft of the present invention breaks through the constraints of the thermocouple measurement range, and effectively reduces the size of the head sensor through integrated measurement, thereby reducing interference with radar electrical performance.

[0084] In order to have a further understanding of the present invention, the total temperature and total pressure measurement method for a high Mach number aircraft of the present invention is described in detail below in conjunction with specific embodiments.

[0085] This embodiment requires measuring the total temperature and total pressure of a certain aircraft.

[0086] Step 1: First, carry out the double-throat nozzle profile design based on the shape of the aircraft head.

[0087] The double-throat nozzle is arranged at the head of the aircraft. First, the initial area ratio of the double Laval throat to the outlet is given as 2.9862, and the opening is selected at a position with low back pressure (x = 200mm). The commercial software FLUENT is used to perform flow field calculations. The simulation results show that the airflow can flow smoothly and reach the speed of sound in both throats, which meets the requirements.

[0088] Step 2: Design the cooling length of the double-throat nozzle according to the cooling range of the aircraft and the total airflow temperature.

[0089] According to the maximum measurable range of the existing tungsten-rhenium thermocouple, the total temperature of the gas needs to be reduced by 200 K. The commercial software FLUENT was used to carry out aerodynamic heat / heat transfer coupling calculations, which confirmed that a cooling length of 150 mm can allow the airflow to be cooled by sufficient radiation heat dissipation, and the total temperature can be reduced to a temperature range that can be directly measured, so the cooling length is set to 150 mm.

[0090] Step three, setting a first opening point before the first throat, opening the hole to bleed air, and measuring the head wall pressure; correcting the head wall pressure to obtain the total pressure before the first throat.

[0091] A hole was opened before the first throat to draw air and measure the head wall pressure. In this example, the first opening point was set at 5 mm behind the head, a 2 mm diameter hole was opened, and air was drawn backward to measure the pressure.

[0092] The commercial software Fluent was used to simulate the flow field, establish the corresponding relationship between the wall pressure 5 mm behind the head and the total pressure before the first throat, and establish the first correction database. The total pressure before the first throat was obtained by interpolation of the measured wall pressure according to the first correction database.

[0093] Step 4: Set a second opening point before the second throat, open the hole to bleed air, and measure the wall pressure; correct the wall pressure to obtain the total pressure before the second throat.

[0094] A hole is opened before the second throat to bleed air and the wall pressure is measured; according to the simulation results of step 1, a hole is opened at a position (x=140mm) in front of the second throat where the flow velocity is lower than 100m / s to bleed air.

[0095] The commercial software FLUENT was used to simulate the flow field, establish the corresponding relationship between the wall pressure at x = 140 mm and the total pressure before the second throat, and establish the second correction database. According to the second correction database, the total pressure before the second throat was obtained by interpolation of the measured wall pressure.

[0096] Step 5: When the airflow before the second throat is cooled to the measurement range of the thermocouple, the total temperature of the airflow before the second throat is measured by using the thermocouple.

[0097] The tungsten-rhenium thermocouple node was arranged at x=140 mm and ensured to be insulated from the inner wall of the nozzle, and the total temperature of the incoming flow before the second throat was measured.

[0098] Step six, based on the sonic flow relationship, obtain the theoretical total enthalpy of the incoming flow according to the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the geometric area of ​​the double nozzle throat.

[0099] In this example, the total enthalpy of the incoming flow h0>2.326×10 6 J / kg, using formula (2) to calculate the theoretical total enthalpy,

[0100]

[0101] Step seven, obtain the theoretical total temperature of the incoming flow based on the theoretical total enthalpy of the incoming flow.

[0102] Formula (3) is used to calculate the theoretical total temperature of the incoming flow.

[0103]

[0104] Step eight, in the arc wind tunnel, simulate the high-speed incoming flow environment, use enthalpy measurement technology to obtain the measured incoming flow total temperature, and calculate the measured incoming flow total enthalpy; obtain the actual area ratio of the double nozzle throat according to the measured incoming flow total enthalpy, so as to obtain the corrected incoming flow total enthalpy calculation formula and the corrected incoming flow total temperature calculation formula.

[0105] The main factor affecting the calculation accuracy in formula (2) is the throat area ratio. However, due to the influence of friction heat transfer loss and boundary layer, the actual throat area is lower than the throat geometric area. Therefore, this step calibrates the value of the throat area ratio by simulating a high-speed incoming flow environment in an arc wind tunnel.

[0106] In this example, the incoming flow environment is simulated in an arc wind tunnel, and the stagnation point heat flow Fay-Riddell formula enthalpy measurement technology of the wind tunnel is used to obtain accurate measured incoming flow total temperature. Then, according to formula (4) and formula (5), the corrected double throat actual area ratio is obtained: Then, the parameters in formula (1) and formula (2) are corrected to obtain the corrected total enthalpy calculation formula (6) of the incoming flow, and then the corrected total temperature calculation formula (7) of the incoming flow is obtained.

[0107] Step nine: After the correction is completed, directly use formula (6) and formula (7) to measure the total temperature and total pressure of the high-speed aircraft.

[0108] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A method for measuring total temperature and total pressure of a high Mach number aircraft, characterized in that: The total temperature and total pressure measurement method for a high Mach number aircraft comprises: According to the shape of the aircraft head, the double-throat nozzle profile design is carried out; Design the cooling length of the double-throat nozzle according to the cooling range of the aircraft and the total airflow temperature; A first opening point is set before the first throat, air is bleeded through the opening, and the head wall pressure is measured; the head wall pressure is corrected to obtain the total pressure before the first throat; A second opening point is set before the second throat, air is introduced through the opening, and the wall pressure is measured; the wall pressure is corrected to obtain the total pressure before the second throat; When the airflow before the second throat is cooled to the measuring range of the temperature sensor, the temperature sensor is used to measure the total temperature of the airflow before the second throat; Based on the sonic flow relationship, the theoretical total enthalpy of the incoming flow is obtained according to the total pressure before the first throat, the total pressure before the second throat, the total temperature of the incoming flow before the second throat, and the geometric area of ​​the double nozzle throat; Obtain theoretical total temperature of incoming flow according to theoretical total enthalpy of incoming flow; In an arc wind tunnel, a high-speed incoming flow environment is simulated, and the measured incoming flow total temperature is obtained by using enthalpy measurement technology, and the measured incoming flow total enthalpy is calculated; the actual area ratio of the double nozzle throat is obtained according to the measured incoming flow total enthalpy, and the theoretical incoming flow total enthalpy calculation formula and the theoretical incoming flow total temperature calculation formula are corrected to obtain the corrected incoming flow total enthalpy calculation formula and the corrected incoming flow total temperature calculation formula; Based on the revised incoming flow total enthalpy calculation formula and the revised incoming flow total temperature calculation formula, the total temperature and total pressure measurement of the high Mach number aircraft is realized.

2. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 1, characterized in that: At the head of the aircraft, the double Laval nozzle profile design is carried out, the flow field calculation is performed, and the area ratio and exit position are adjusted.

3. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 1, characterized in that: The first opening point should be less than 10mm from the front edge of the head.

4. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 1, characterized in that: By carrying out flow field simulation, the corresponding relationship between the head wall pressure at the first opening point and the total pressure before the first throat is established, and a first correction database is established; the total pressure before the first throat is obtained according to the first correction database.

5. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 4, characterized in that: By carrying out flow field simulation, the corresponding relationship between the wall pressure at the second opening point and the total pressure before the second throat is established, and a second correction database is established; the total pressure before the second throat is obtained according to the second correction database.

6. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 1, characterized in that: The center of the temperature sensor node is located in the same plane as the center of the second opening point and is insulated from the inner wall of the nozzle.

7. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 5, characterized in that: according to Calculate the theoretical total enthalpy of the incoming flow, where h0 is the theoretical total enthalpy of the incoming flow, A1 and A2 are the geometric areas of the first throat and the second throat respectively, is the total pressure before the first throat, is the total pressure before the second throat, and is the total temperature of the incoming flow before the second throat.

8. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 7, characterized in that: according to Calculate the theoretical total temperature of the incoming flow, where c p is the specific heat of air and T is the temperature.

9. The total temperature and total pressure measurement method for a high Mach number aircraft according to claim 8, characterized in that: The actual area ratio of the double nozzle throat is based on Get, where is the actual area ratio of the double nozzle throats, is the measured total enthalpy of the incoming flow, according to Get, It is the measured total temperature of the incoming flow.

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